This disclosure relates to a supplemental radial cooling passage for an airfoil.
Turbine blades are utilized in gas turbine engines. As known, a turbine blade typically includes a platform having a root on one side and an airfoil extending from the platform opposite the root. The root is secured to a turbine rotor. Cooling circuits are formed within the airfoil to circulate cooling fluid, such as air. Typically, multiple relatively large cooling channels extend radially from the root toward a tip of the airfoil. Air flows through the channels and cools the airfoil, which is relatively hot during operation of the gas turbine engine.
Some advanced cooling designs use one or more radial cooling passages that extend from the root toward the tip. Typically, the cooling passages are arranged between the cooling channels and an exterior surface of the airfoil. The cooling passages provide extremely high convective cooling.
The Assignee of the present disclosure has discovered that in some cooling designs the airfoil is overcooled at the base of the airfoil near the platform. It is believed that strong secondary flows, particularly on the suction side, force the migration of relatively cool fluid off the end wall and onto the suction side of the blade. This results in relatively low external gas temperatures. Internally, the coolant temperature is relatively cool as it has just entered the blade. The high heat transfer coefficients provided by the cooling passage in this region are undesirable as it causes overcooling of the external surface and premature heating of the coolant air.
Tapered radial cooling passages have been used. However, in one arrangement, the wall adjacent to the suction side exterior surface is tapered as it extends towards the tip. This configuration undesirably results in increased cooling near the platform as compared to near the tip due to the larger convection surface near the platform.
In another arrangement in which the cross-sectional area of the cooling passage remains relatively constant cooling fluid, Mach numbers also remain relatively constant resulting in uniform heat transfer rates within the passage. Coolant fluid entering the airfoil at low temperature and increases in temperature as it moves through the cooling passage. External three-dimensional flows and non-uniform gas temperature profiles cause temperatures and heat transfer rates to be typically lower near the inner and outer radii of the airfoil. This external heat load, combined with the cool coolant fluid near the inlet to the airfoil cause the external surface to be overcooled.
What is needed is a radial cooling passage that provides desired cooling of the airfoil.